FIG. 1 shows a cross-section through a portion of a turbine engine. A turbine engine 10 can generally include a compressor section 12, a combustor section 14 and a turbine section 16. A centrally disposed rotor 18 can extend through the three sections.
Generally, the combustor section 14 is enclosed within a casing 20 that can form a chamber 22, together with the aft end of the compressor casing 24 and a housing 26 that surrounds a portion of the rotor 18. A plurality of combustors 28 and ducts 30 can be provided within the chamber 22, such as in an annular array about the rotor 18. Each duct 30 can connect one of the combustors 28 to the turbine section 16.
The turbine section 16 can include an outer casing 32 which encloses alternating rows of stationary airfoils 34 (commonly referred to as vanes) and rotating airfoils 36 (commonly referred to as blades). Each row of blades can include a plurality of airfoils 36 attached to a disc 38 provided on the rotor 18. The rotor 18 can include a plurality of axially-spaced discs 38. The blades 36 can extend radially outward from the discs 38 and terminate in a region known as the blade tip 40.
Each row of vanes can be formed by attaching a plurality of airfoils 34 to the stationary support structure in the turbine section 16. For instance, the airfoils 34 can be hosted by a vane carrier 42 that is attached to the outer casing 32. The vanes 34 can extend radially inward from the vane carrier 42 or other stationary support structure to which they are attached.
In operation, the compressor section 12 can induct ambient air and can compress it. The compressed air 44 from the compressor section 12 can enter the chamber 22 and can then be distributed to each of the combustors 28. In the combustors 28, the compressed air can be mixed with the fuel introduced through a fuel nozzle 46. The air-fuel mixture can be burned, thereby forming a hot working gas 48. The hot gas 48 can flow through the ducts 30 and then through the rows of stationary airfoils 34 and rotating airfoils 36 in the turbine section 16, where the gas 48 can expand and generate power that can drive the rotor 18. The expanded gas 50 can then be exhausted from the turbine 16.
It should be noted that each row of blades 36 is surrounded by the stationary support structure of the turbine, which can be the outer casing 32, the vane carrier 42 or a ring seal (not shown). The space between the blade tips 40 and the neighboring stationary structure is referred to as the blade tip clearance C. During engine operation, gas leakage can occur through the blade tip clearances C, resulting in measurable engine performance decreases in power and efficiency.
While small blade tip clearances C are desired to minimize gas leakage, it is critical to maintain a clearance C between the rotating turbine components (blades 36, rotor 18, and discs 38) and the stationary turbine components (vanes 34, outer casing 32, vane carriers 42 and ring seals) at all times. Rubbing of any of the rotating and stationary components can lead to substantial component damage, performance degradation, and extended outages.
However, during transient conditions such as during engine startup or part load operation, it can be difficult to ensure that adequate blade tip clearances C are maintained because the rotating parts and the stationary parts thermally expand at different rates. For instance, in a cold start situation, the rate of thermal expansion of the thermal stationary support structure is at least initially less than the rate of thermal expansion of the rotating turbine components due to the relatively larger size and thickness of the stationary support structure. As a result, the blade tip clearances C can actually decrease because the rotating components expand radially outward faster than the stationary support structure, raising concerns of blade tip rubbing.
To avoid blade tip rubbing, large tip clearances are initially provided so that minimum blade tip clearances C are maintained at known pinch points, that is, during operational conditions where the clearances C would otherwise be expected to be the smallest (hot restart, spin cool, etc.). However, because the minimum blade tip clearances C are sized for these pinch point conditions, the clearances C eventually become overly large as the rate of thermal expansion of the rotating components slows or substantially stops while the stationary support structure continues to grow radially outward. Such oversized clearances C can occur as the engine approaches or attains steady state operation, such as at base load. Consequently, engine power and efficiency can be reduced.
Thus, there is a need for a system that can improve engine performance by minimizing turbine tip clearances at desired engine operating conditions.